The application described herein relates generally to gas turbine engines, and more specifically to a turbine frame cooling systems for use in a gas turbine engine.
Gas turbine engines typically include an inlet, a fan, low and high pressure compressors, a combustor, and low and high pressure turbines. The compressors compress air which is channeled to the combustor where it is mixed with fuel. The mixture is then ignited for generating hot combustion gases. The combustion gases are channeled to the turbine(s) which extracts energy from the combustion gases for powering the compressor(s), as well as producing useful work to propel an aircraft in flight or to power a load, such as an electrical generator.
During engine operation, significant heat is produced which raises the temperature of the frames that surround various engine components. Specifically, at least one known frame includes radial support struts which project across an annular flowpath to interconnect the inner and outer frame members. Since the temperature of the motive fluid flowing through the flowpath changes very rapidly during transient engine operation, substantial thermal stresses can be created in the rigid frame assemblies when the struts are allowed to heat up and cool down at rates differing substantially from those of the inner and outer frame members. This is particularly true with respect to the turbine frame assembly since the exhaust gases which surround the turbine frame assembly are subject to the most rapid and greatest changes in operating temperatures and resulting thermal stresses. At least some known cooling systems use compressor air or bore air to purge and cooling the frame to reduce the thermal gradients. However, the use of compressor air or bore flow may result in less efficient engine cycle.